Turbine engine equipped with thrust reverser

ABSTRACT

A thrust reverser system for a turbine engine of the type having a central relatively hot gas generator nozzle conducting a core flow and a relatively cold duct conducting a fan flow and surrounding the hot flow, a pair of thrust reverser door, each of the doors being pivotally mounted on an axis which is substantially diametrically positioned with respect to the exhaust nozzle of the engine so as to pivot between a stowed position in which the doors are out of the direct path of exhaust from the engine and a deployed position in which the doors are in the path of the engine exhaust for deflecting the exhaust and creating a braking thrust, a flow modifying arrangement acting on the hot gas generator nozzle for simultaneously decreasing the total pressure of the core flow and increasing the total pressure of the fan flow by increasing the area of the hot gas generator nozzle at the discharge end thereof and moving the boundary between the hot gas and cold gas flows radially outwardly.

CROSS-REFERENCE TO RELATED APPLICATION

This application is a continuation in part of my application Ser. No.14,550 filed Feb. 8, 1993, now U.S. Pat. No. 5,372,006.

This invention relates to a turbine engine equipped with a thrustreverser, especially the type used on aircraft. More particularly, theinvention relates to a turbine engine of the double flow type providedwith a thrust reverser, and providing for area variation of the hotcentral nozzle.

BACKGROUND AND OBJECTS OF THE INVENTION

In order to reduce the landing distance of a jet engine poweredaircraft, as well as to increase the margin of safety when the aircraftis landing on a wet or icy runway, aircraft jet engines are providedwith thrust reversers in order to provide a braking thrust for theaircraft. Typically, such thrust reversers are formed by thrust reverser"doors" which are capable of pivoting between two positions about anaxis which is transverse and substantially diametrical with respect tothe jet of the engine.

The first position finds the doors in a stowed position, out of thedirect path of the exhaust blast of the engine. In this position, thedoors form the exhaust nozzle of the gas turbine engine so that thethrust of the engine is directly rearward, thereby producing the forwardthrust of the aircraft. In the second position, the doors are pivotedabout the pivot axis to a transverse, blast deflecting or deployedposition, to intercept and redirect the jet blast and produce thebraking thrust for the aircraft when needed.

The jet engines which utilize such thrust reversers are typically of thedouble jet type of gas turbine engines. Such engines comprise a centralgenerator which emits a jet of hot gas, and an annular by-pass conduitsurrounding the central generator, and through which passes a jet ofrelatively cold gas. In practice, the central or hot jet emits gas flowat a high temperature on the order of 500° C. to 600° C., while theby-pass flow is at a temperature considered relatively cold, i.e. on theorder of 100° C.

When a thrust reverser is used with such double flow engines, the thrustreverser re-directs in a forward direction the hot and cold flows, inorder to produce the braking thrust. Since the hot, central flowimpinges on the thrust reverser doors, they must be made of heatresistant material. In general, the material used for the thrustreverser doors is either steel or a titanium alloy. However, steel is arelatively heavy metal, and while titanium alloys are much lighter, theyare considerably more expensive. Aluminum is a very desirable materialfor the thrust reverser doors, but because of the high temperature ofthe hot flow, aluminum cannot be used.

In the past, attempts to overcome the drawbacks of conventional doubleflow engines have been made, for example in U.S. Pat. Nos. 4,362,015 and4,581,890. For example, it has been proposed to provide an obstructionin the nature of flaps for obstructing a portion of the discharge outletarea of the exit end of the by-pass conduit when the thrust reverser isdeployed. According to U.S. Pat. No. 4,362,015, the flaps restrict theflow of cold gas in by-pass conduit. In U.S. Pat. No. 4,581,890, theflaps are said to briefly reduce the efficiency of the flow mixing andcausing the cold by-pass flow to contact the structure of the reverser,which was thereby maintained at a lower temperature. More precisely,this prior system increased (in thrust-reversing mode) the value of theratio of the total pressure of the by-pass flow (i.e. the cold flow) tothe total pressure of the hot flow, so that only the cold flow impingeson the reverser doors, thereby enabling the use of aluminum alloy doors.

However, it has been found that when such systems are applied to turbineengines that significant drawbacks arise. In essence, tests have shownthat the engine rating must be kept reduced in the reverse mode in orderto avoid surging the engine and raising the skin temperature of thethrust reverser doors. This is particularly true when the central nozzleof the gas generator is of the mixer type. The limitations found in theprior art systems are primarily due to the means used to control thevalue of the ratio of the fan pressure to the core pressure P_(T)(fan)/P_(T)(core). The prior art systems sought to increase this ratio bydirectly changing the flow characteristics of the fan and increase thefan pressure by mechanically restricting the fan flow by-pass area inthe reverse mode, while only indirectly affecting the core flow.

As described in U.S. Pat. Nos. 4,362,015 and 4,581,890, these systemsdirectly act on the total pressure of the fan flow so as to increase it,thereby increasing the ratio. This is done by using flaps in the coldflow to increase the pressure of the cold flow. However, the priorsystems have no significant direct effect, and only insignificantindirect effect, on the value of the total pressure of the core flow.Ordinarily, the ratio would be on the order of 0.9-1.0, and through theuse of the flap techniques disclosed in these prior patents, the ratiowas increased to a value at most on the order of 1.2. Experience hasshown that in some cases it has been necessary to install more than two,for example 3 or 4, flaps in the fan stream to adequately increase thepressure ratio of P_(T)(fan) /P_(T)(core) in reverse. As a result, thevalue of the total pressure ratio (fan/core) is increased in reverse,but mainly by the increase of the P_(T)(fan). However, this is notdesirable for the engine, as it increases the risk of surging of theengine. It is significant that in accomplishing the increase in theratio, only the by-pass flow was acted upon, and this by the use of twoor more flaps obturating the fan stream at the outlet end of the bypassconduit.

Accordingly, a primary object of the present invention is to overcomethe drawbacks of the prior art thrust reverser systems.

Another object of this invention is to provide a thrust reverser whichcan use lightweight, inexpensive alloys such as aluminum, or compositematerials, for the thrust reverser doors.

A further object of the invention is to provide an improved thrustreverser which operates with a significant increase in the ratio of thepressure of the fan flow to the pressure of the core flow when thethrust reverser is deployed.

Yet another object of the invention is to provide an improved thrustreverser system which increases the ratio of the total fan flow pressureto the total core flow pressure by directly acting upon the core flow.

Still another object of the invention is to provide an improved thrustreverser system which increases the ratio of the total fan flow pressureto the total core flow pressure by increasing the area of the hot gasnozzle while decreasing the area of the fan duct.

Still a further object of the invention is to provide an improved doubleflow jet engine which enables the use of a thrust reverser made oflightweight materials, by providing a nozzle which enables a mixing ofthe hot and cold flows with simultaneous decreasing of the totalpressure of the hot flow and increasing the total pressure of the coldflow.

Another object of this invention is to provide a process for improvingthe efficiency of a double flow turbine engine by adjusting the mixingof the cold and hot flows, thereby simultaneously increasing the totalpressure of the cold flow and decreasing the total pressure of the hotflow, by acting directly and primarily on the hot flow stream.

These and other objects and advantages of the invention will becomeapparent from a detailed consideration of the following description andclaims, when taken together with the accompanying drawings.

DESCRIPTION OF THE INVENTION

According to the present invention, the total pressure of each flowstream, both core and fan, is changed. However, an increase in thepressure of the cold flow is generally only possible with very narrowlimits, dictated by the surge margin of the low pressure compressor.Therefore, the total pressure of the fan flow is increased to a slightextent, while the total pressure of the core flow is decreased to agreater extent, so that the ratio P_(T)(fan) /P_(T)(core) issignificantly altered. In this manner the risk of surging the lowpressure compressor of the engine is minimized. This, then, isaccomplished by varying the exhaust area of the central generatornozzle, in the reverse thrust mode. In reverse, the exhaust area of thecentral nozzle is increased in conjunction with the deployment of thereverser doors. In the forward thrust mode, during take-off and climb,the exhaust area of the central nozzle is increased in conjunction withthe throat area variation of the final exhaust nozzle.

This invention not only allows a significant weight reduction of thethrust reverser doors (since lightweight aluminum alloy or compositematerial may be used) but also allows further improvements of engineperformance when the engine is fitted with a thrust reverser whichintegrates a variable exhaust area nozzle as described in U.S. Pat. No.5,181,676, or with a non-reversing variable exhaust area nozzle asdescribed in my co-pending application Ser. No. 07/741,647, thespecifications of both of which are incorporated herein by reference.

DESCRIPTION OF THE DRAWINGS

This invention will be described in greater detail by way of referenceto the accompanying drawings, which show by way of non-limitingexamples, certain preferred features and embodiments of this invention,and in which:

FIG. 1 is a cross-sectional schematic view of the rear portion of aturbine engine fitted with a central mixing nozzle;

FIG. 2 is a cross-sectional schematic view in the outlet plane of theturbine mixer nozzle along lines 2--2 of FIG. 1 and viewed in thedirection of the arrows;

FIG. 3 is a view similar to FIG. 2 with the exhaust area of the mixernozzle increased;

FIG. 4 is a magnified cross-sectional of a portion of the mixer nozzleshowing the lobe in its retracted and deployed position;

FIG. 5 is a view similar to FIG. 4, showing an alternate embodiment ofthe invention;

FIG. 6 is a cross-sectional view along lines 6--6 of FIG. 4 and viewedin the direction of the arrows;

FIG. 7 is a cross-sectional view along lines 7--7 of FIG. 4 and viewedin the direction of the arrows;

FIGS. 8 and 8a are views respectively similar to FIGS. 4 and 6 ofanother alternative embodiment;

FIG. 9 is a cross-sectional schematic view of the rear portion of aturbine engine fitted with a central non-mixing nozzle;

FIG. 10 is a cross-sectional view along lines 10--10 of FIG. 9 andviewed in the direction of the arrows; and

FIG. 11 is a view similar to FIG. 9 showing another embodiment of theinvention;

FIG. 12 is a cross-sectional schematic view of the rear portion of aturbine engine fitted for aerodynamic flow modification, in the stowedconfiguration; and

FIG. 13 is view similar to FIG. 12 showing the modified flow embodimentin the deployed configuration.

DETAILED DESCRIPTION OF PREFERRED EMBODIMENTS

Referring first to FIG. 1, the rear portion of a turbine engine 10 isshown, and is provided with a central hot gas generator 12 of the mixingtype. This hot gas generator 12 is surrounded by a peripheral fixedstructure which is formed by an inner skin 14 and an outer skin 16. Theannular conduit 18 is formed between the inner skin 14 and the hot gasgenerator 12, and provides a passage for the by-pass flow of fan air,which is at relatively cold temperatures. On the downstream end of theinner skin 14 is installed a thrust reverser which is of the targettype, formed of two doors 20. Preferably this thrust reverser is of thetype described in U.S. Pat. Nos. 5,176,340 and 5,181,676.

As shown in FIG. 2, the primary exhaust nozzle 12 is formed by aplurality of lobes 22, through which the hot gases generated by theturbine engine 10 exit. Between each lobe there is a channel 24, and theby-pass flow (again at relatively low temperature) flows through thesechannels. The lobes 22 and channels 24 are adapted to guide respectivelythe hot gases (also called the core flow of the engine) and the coldgases (called the fan flow) toward each other, so that these two gasflow paths are caused to mix downstream of the primary exhaust nozzle12.

As shown in FIGS. 1 through 3, and FIGS. 6 and 7, the primary nozzle 12has the capability of having its exit area increased in different ways.In the embodiments shown, this increase in the area is achievedgeometrically by means of movable lobes 26 which are hinged on theirupstream end. In their retracted positions (FIG. 2), the exit area ofthe hot gases through the plurality of lobes 22 is at a minimum, and thethrust reverser doors 20 are in the stowed configuration (the solid lineposition in FIG. 1). In their fully extended positions (FIG. 3), thelobes 26 increase the exit area of the primary nozzle 12 through whichflow the hot gases generated by the turbine engine. This results in achanging of substantially the entire boundary between the hot and coldstreams, moving that boundary radially outwardly. Although FIG. 2 showstwo movable members of three lobes each, it will be understood that thenumber of movable lobes can vary as a function of the percentage ofincreased area which is targeted for the exhaust area of the primarymixer nozzle. This exhaust area may also be increased aerodynamically byestablishing a divergence following the contour of the lobe(s) in aportion of the trailing edge of the mixer nozzle. This could be done bymeans of a retractable divergent extension member(s) normally housed outof the flow streams, typically in the structure surrounding the hot gasgenerator, when the thrust reverser is in the stowed configuration, anddeployed to form a divergent extension(s) of the core nozzle, thereforechanging the boundary between the hot and cold streams, and moving itradially outwardly when the thrust reverser is deployed.

In the geometric variation embodiment shown in FIGS. 1-3, when thethrust reverser is commanded to deploy, the movable lobes of the primarymixing nozzle are commanded to their fully extended position. As aresult of the movable lobes 26 being positioned in their fully extendedposition, the exhaust area of the primary exhaust nozzle is increased,while the by-pass area of the annular conduit is decreased, because ofthe penetration of the movable lobes into the annular conduit 18, theboundary between the hot and cold flow is changed, and moved radiallyoutwardly. Because of the increased exhaust area of the primary mixernozzle, the total pressure of the hot gases is decreased, while thetotal pressure of the by-pass flow is increased because of the decreaseof the by-pass area. The core area increase being greater in absolutevalue than the corresponding decrease of the fan by-pass area, thepressure of the core stream decreases more than the pressure increase inthe fan flow. Therefor, the ratio of the total pressure of the by-passfan flow to the total pressure of the core flow is significantlyincreased when the reverser doors are deployed. Also, the penetration ofthe movable lobes into the fan stream serves as an aerodynamic deflectorto the fan stream without blocking of the fan stream. In the reversingstate, by significantly decreasing the total pressure of the cold flowfor a small increase of the total pressure of the fan flow, ensures thatthe by-pass or cold flow will have more energy and the ability tocontain the core flow or hot gases generated by the turbine engine,thereby preventing the hot gases from impinging on the reverserstructure. Therefor, the by-pass flow serves as a thermal shield for thereverser structure, allowing the use of conventional lightweightmaterial such as an aluminum alloy or a composite material such asgraphite for its construction. Overall weight of the thrust reverserassembly is greatly reduced, as well as the manufacturing costs relatedto the thrust reverser.

With reference to FIGS. 6 and 7, one embodiment of the movable lobes isshown in greater detail. Each movable lobe is of the same shape of thefixed lobe that it covers, i.e. is composed of two flat walls 32 joinedby a radius 34. This configuration, while presenting a geometricincrease of the area of the core flow, also presents an aerodynamicallyefficient surface to the fan flow. Each flat wall 32 incorporates in itslower portion a metallic seal 36 which ensures both hot gas tightnessbetween the walls of the movable lobes and the adjacent lobe of thefixed wall. In the movable multi-lobe arrangement as shown in FIGS. 2,3, 6 and 7, the central movable lobe is the guide, while the twoadjacent lobes are followers. Of course without departing from thespirit of the invention, the two outer lobes may be guides, and themiddle lobe a follower.

To achieve the guide/follower operation of the lobes, the guide lobe isprovided on each of its side walls 32 with a double shoulder 38, 40,while the follower lobes have a single shoulder 42 which engages theguide shoulders 38, 40. When the lobes have reached their fully openedposition, on their downstream end the shoulder of the follower lobe maydisengage the double shoulder of the guide lobe as shown in FIGS. 3 andthe dotted line position of FIG. 6, but the guiding characteristic ofthe lobes is still ensured as shown in FIG. 7 by the shoulders 38, 40and 42 which in their upstream portion never become disengaged.

Still with reference to FIGS. 6 and 7, it is apparent that the shoulders38, 40 and 42 also have the added function of changing the area ofchannel 8 through which the fan air (or cold flow) flows. The by-passarea variation is also achieved by the penetration of the movable lobesinto the by-pass conduit, or other similar variations, moving the lobesor shoulders between the lobes. The shoulders 38, 40 and 42 serve alsoas aerodynamic deflectors to the fan flow.

When the movable lobes are positioned to their fully extended positionas seen in FIGS. 3 and 6, they restrict the area of the by-pass conduit,guide and deflect the fan flow, and increase the area of the mixernozzle. As a result, the ratio of the total pressure of the by-pass flowover the total pressure of the core flow is increased, which in turn,and together with the deflection of the fan flow, prevent the core flowfrom impinging on the thrust reverser structure when it is deployed, asdiscussed above.

Another feature of the present invention allows for increasing the areaof the central mixing nozzle in the forward thrust mode, particularlyduring take-off under high outside air temperatures. In such cases, itis beneficial that the final exhaust nozzle be either a thrust reverserwhich incorporates a variable exhaust area nozzle system as described inU.S. Pat. No. 5,181,676, or a non-reversing variable exhaust nozzle asdescribed in copending application Ser. No. 07/741,647. If the rear exitof the nacelle is equipped with a thrust reverser which integrates avariable exhaust area nozzle, such as the type described in U.S. Pat.No. 5,181,676, then in the forward thrust mode, the exhaust area of themixer nozzle can be adjusted together with the exhaust area of thereverser. This results in increased take-off thrust under high outsideair temperature. The increase of the exit area of the mixer nozzle wouldalso have a direct effect on the operating temperature of the engine, bydecreasing it. This temperature decrease happens at the most criticalengine power setting, i.e. at maximum engine thrust which are the mostadverse engine operating conditions. Any decrease of the maximum engineoperating temperature has a direct effect on the engine life,significantly increasing the life.

In the thrust reversing mode, the exhaust area of the mixer nozzle isincreased as described previously in order to increase the value of theratio of the total pressure by-pass flow over core flow, so that thecore flow never impinges on the reverser structure, allowing use ofconventional aluminum alloy or composite materials. If the rear end ofthe nacelle is equipped with a non-reversing variable exhaust nozzle asdescribed in application Ser. No. 07/741,647, then in the forward thrustmode, which is the only mode of operation in that case, the exhaust areaof the mixer nozzle is adjusted together with the exhaust area of thefinal variable exhaust nozzle. As a consequence, take off thrust underhigh outside air temperature conditions is increased and maximum engineoperating temperature is decreased, thus considerably improving theengine life.

Referring to FIG. 4, the control system for the angular positioning ofthe movable lobes is shown. This system comprises at least one actuator50 connected to linkage members 52 and 54. Each movable lobe is hingedon each side of its respective wall 11 (FIG. 6) on its upstream endthrough hinge fitting 56. These hinge fittings 56 are supported by hingesupport 58 attached to the bottom fixed wall of the channel 24. The arm54 which controls the angular positioning of the guide lobe is hinged onthe same axis of rotation as the guide lobe. One end 58 of this arm 54is attached to the movable lobe 30, while the other end 60 is connectedto link 52. When the movable lobe 30 is in position A (small exit areaof the mixer nozzle) then the linkage members 52 and 54 are self-lockedin that position because of the over center relationship of link 54 andthe rod 62 of the actuator 50.

When the actuator 50 is pressurized in the direction of arrow 64, therod 62 retracts causing its end 66 to follow the guide 68, forcing thelobe 30 to close a little more until connecting link 52 has overcome theovercenter position of the linkage arrangement. The guide lobe can thenrotate freely, driven by its control actuator 50 to reach position B.The follower lobes as described previously reach a similar positionsince they are driven by the guide lobe.

If the system is operating in the reverser mode, then the reverser door20 starts the deployment sequence represented by the dotted lines 20a.

If the system is operating in the forward thrust mode, then the areaincrease of the mixer nozzle is achieved in conjunction with the areaincrease of the final exhaust nozzle as in U.S. Pat. No. 5,181,676 orapplication Ser. No. 07/741,647. In such a case, it would be beneficialthat the control system of the guide lobe use a screw and nut so that itcan be positioned to any intermediate angular position required toachieve targeted performance.

FIG. 5 shows a variation of the control of the angular position of theguide lobe. The control system is completely contained in the box formedby the inner skin 14 and the outer skin 16 of the fixed structuresurrounding the hot gas generator. This embodiment has the added benefitof completely placing the control system in a "cold" environment, i.e.out of the hot gas flow. A profiled link 70, to reduce drag, has one end72 hinged on a fitting 74 attached to the guide lobe 30, while its otherend 76 is hinged inside the box formed by the inner skin 14 and outerskin 16, to a V-shaped crank 78. The crank 78 is hinged on a supportfitting 80 attached to the outer skin 16 of the fixed structuresurrounding the hot gas generator.

The other arm of the V-shaped crank 78 is connected to a link 82 whichin turn is connected to a V-shaped lever 84 hinged on the close-outstructure 86 of the inner skin 14 and outer skin 16. The lever 84 hasone of its arms always remaining outside of the box formed by the skins14, 16, and the close out 86. Still with reference to FIG. 5, theangular position of the lever 84, and therefor the angular position ofthe guide lobe 30, is directly controlled by the reverser door 20. Whenthe reverser door starts its deployment sequence, the wheel 88 liberatesthe lever 84 which, through the combined action of spring 90, thelinkage arrangement and the pressure acting in the guide and followerlobes, causes the complete system controlling the angular position ofthe guide lobes to rotate to the position shown by dotted lines. Whenthe reverser door is moving from its deployed position to its stowedposition, then when wheel 88 comes in contact with the V-shaped lever84, forcing it to return to its original (solid line) position, wheel 88contacts the V-shaped lever 84 forcing it, and the complete system, toreturn to its original (solid line) position.

It will be noted that while link 70 penetrates the inner skin 14, fluidtightness is still ensured through seal 92 and cover 94, the cover 94being part of link 70. A leaf spring seal 96 ensures hot gas fluidtightness for any angular position of the movable lobes.

The system described in FIG. 5 can only operate with a thrust reverserin the reverse mode. If the system is required to be operated in theforward thrust mode as well, then an actuator controlling directly theangular position of the V-shaped lever 78 would need to be provided. Insuch case, link 82, lever 84 and spring 90 are deleted.

FIG. 8 shows a variation of the mixer nozzle. Three fixed lobes 100 areprovided with a plurality of perforations 102 on their upper portionsand lower portions. These perforations 102 are covered by the movableguide and follower lobes when they are retracted. When the guide andfollower lobes are controlled to their deployed position through one ofthe control systems previously described, the internal pressure P actinginside the mixer nozzle decreases because of the opening of cavities 104and 106. At the same time, the by-pass area 18 of the cold flow conduitis decreased because of the penetration of the movable lobes in thatconduit. The total pressure of the cold flow is therefor increased. Theratio of the total pressure of the cold flow to the total pressure ofthe hot flow is therefore increased in the reverser mode, ensuring thatthe hot flow will not impinge on the reverser structure.

Referring to FIG. 9, 10 and 11, the central generator is provided with afrustoconical nozzle (a non-mixing nozzle). With this type of primarynozzle, the hot gases and the cold gases are no longer forced to mix. Asseen in FIG. 9 and 10, the wall 110 of the conical nozzle is cut at itsupper and lower parts. The structural integrity of the nozzle isre-ensured by the housing 112. Two flaps 114 and 116 are installed. Flap114 deploys inside the housing 112, while flap 116 deploys outside ofthe housing 112. Flap 114 is hinged at its upstream end on wall 110;when in the non-deployed position (solid line position in FIGS. 9 and10, it ensures the smooth profile continuity of the inner contour of thewall 110. Metallic seals 118, 120 and 122 are provided at the peripheryof the flap 114 to ensure fluid tightness with respect to hot gases.Flap 116 is hinged at its downstream end on housing 112. Although thehinge point of flap 116 is shown as being outside of housing 112, it canbe located inside of housing 112 without departing from the spirit ofthe invention.

When in the non-deployed position (solid line position of FIG. 9), flap114 ensures the profile continuity of housing 112. Flaps 114 and 116 arelinked by a linkage 124. The angular position of flap 116 is controlledby a profiled link 126 in a similar way as the control system shown inFIG. 5. The main difference from FIG. 5 is that when flap 114 starts itsdeployment driven by the link 126, then it drives with it flap 116because of the link 124 which can move through the slot 128 in housing112. When both flaps are deployed, the total pressure of the hot gasesis decreased because of the increase of area of the nozzle, and thetotal pressure of the cold gases is increased because of the decrease ofthe by-pass area resulting from the deployment of the flap 116 in theby-pass conduit 18. As a result of the above, the ratio of the totalpressure of the by-pass flow to the total pressure of the core flow isincreased, preventing the hot gases from impinging on the structure ofthe reverser door. The by-pass flow or cold gas stream serves as athermal shield to the reverser structure. Again, the system iscompletely controlled by the reverser door. This exhaust area may alsobe increased aerodynamically by establishing a divergence in a portionof the trailing edge of the nozzle. This could be done by means of aretractable divergent extension member(s) normally housed out of theflow streams, typically in the structure surrounding the hot gasgenerator, when the thrust reverser is in the stowed configuration, anddeployed to form a divergent extension(s) of the core nozzle, thereforechanging the boundary between the hot and cold streams, and moving thatboundary radially outwardly when the thrust reverser is deployed.

If it is desired to adjust, in the forward thrust mode, the exhaust areaof the frustoconical hot nozzle, then it becomes necessary toindependently control the flaps 114 and 116, as shown in FIG. 11. Flap114 is controlled by a similar system as the one described in FIG. 4.During the angular rotation of flap 114 from its stowed position (solidline) to its deployed position (dotted lines), the link 126 is allowedto translate on the connecting pin 130 of flap 116 because of the slot132 in the link 126. During that motion of flap 116, flap 116 remains inthe stowed configuration (solid lines). Flap 114 can be returned throughits control actuator (not shown) to its stowed position (solid lines).Flap 116 can be deployed through a linkage arrangement symbolized byarrow 134 (similar to the linkage shown in FIG. 9). Flap 114 needs to bepositioned to its deployed position before flap 116 starts itsdeployment.

FIGS. 12 and 13 show an embodiment for aerodynamically increasing theexhaust area. In this embodiment, a retractable member 150 (of which atleast two would be provided) is stowed in the jet pipe 152. When thethrust reverser doors 154 are deployed, the member 150 pivots toward thecore nozzle 156 creating a divergent core flow area as seen in FIG. 13.

While this invention has been described as having certain preferredfeatures and embodiments, it will be understood that it is capable ofstill further variation and modification without departing from thespirit of the invention, and this application is intended to cover anyand all variations, modifications and adaptations as may fall within thespirit of the invention and the scope of the appended claims.

I claim:
 1. A thrust reverser system for a turbine engine of the typehaving on a longitudinal axis a central hot gas generator nozzleconducting a relatively hot core flow and a fan duct conducting arelatively cold flow surrounding the hot flow, comprising a pair ofthrust reverser door members, each of said door members being pivotallymounted on a pivot axis which is substantially diametrically positionedwith respect to the exhaust of the engine so as to pivot between astowed position in which said door members are out of the direct path ofexhaust from the engine and a deployed position in which said doormembers are in the path of the engine exhaust for deflecting the exhaustand creating a braking thrust, aerodynamic means in the outer wall ofsaid fan duct for simultaneously decreasing the total pressure of saidcore flow and increasing the total pressure of said fan flow at thedischarge end thereof.
 2. A thrust reverser system as in claim 1 andwherein said aerodynamic means comprises a retractable member stowed insaid fan duct for diverting at least a portion of said fan flow.
 3. Athrust reverser system as in claim 2 and wherein said retractable memberforms a divergent extension of said nozzle while diverting said fanflow.
 4. A thrust reverser system as in claim 1 and wherein saidaerodynamic means comprises a retractable member stowed in said fan ductfor aerodynamically expanding at least a portion of the exhaust area ofthe hot gas generator nozzle.
 5. A thrust reverser system as in claim 4and wherein said retractable member forms a divergent extension at thetrailing edge of said hot core nozzle while diverting said fan flow. 6.A process for modifying the flow of gas through a turbine engine of thetype having a central relatively hot gas generator nozzle conducting acore flow and an annular duct for conducting a relatively cold fan flowsurrounding the hot flow for improving the performance of said engine,said process comprising providing diversion means in the outer wall ofsaid annular duct and diverting at least a portion of the fan flow fordecreasing the total pressure of said hot core flow and simultaneouslyincreasing the total pressure of said fan flow and moving the boundarybetween said core flow and said fan flow radially outwardly within saidduct.
 7. A process as in claim 6 and wherein said diverting stepcomprises pivoting a flap into said cold fan flow and simultaneouslyforming with said flap a divergent extension of said hot gas generatornozzle.